Referring to FIGS. 1–2, the turbine section 10 of a turbine engine includes a rotor 12 having a longitudinal axis 14. A plurality of discs 16 (only one of which is shown) are provided on the rotor 12; the discs 16 are axially spaced from each other. A plurality of blades 18 (only one of which is shown) are mounted on each disc 16 to form a row of blades 18. The blades 18 are arrayed about the periphery of the disc 16 and extend radially outward therefrom.
Along the axial direction of the turbine 10, rows of blades 18 alternate with rows of stationary airfoils or vanes 20. Unlike the blades 18, the vanes 20 are attached at one end to a blade ring or casing 21 and extend radially inward therefrom to a radially inner end, referred to as an inner shroud 22. Any of a number of devices can be attached to the inner shroud 22. In the first row of vanes, for example, a pre-swirler 24 can extend from the inner shroud 22. Because the rows of stationary airfoils 20 and the rows of rotating airfoils 18 are spaced from each other, there are axial gaps 26 between these components.
In general, the turbine section 10 includes a radially outer region 28 and a radially inner region 30. Hot gases from the combustor section (not shown) of the engine are directed toward the radially outer region 28 of the turbine 10, which includes the alternating rows of stationary airfoils 20 and rotating airfoils 20. These components can withstand the high temperature of the combustion gases. In contrast, components in the radially inner region 30, such as the discs 16, can fail if exposed to the hot combustion gases. Accordingly, these components must be protected from the hot combustion gases. However, protecting the discs 16 and other components in the radially inner region 30 can be difficult because the axial gaps 26 provide a leak path for the hot gases to penetrate the radially inner region 30 of the turbine 10. While some leakage may be inevitable, there are various techniques for minimizing the amount of leakage or diminishing the severe consequences of such infiltration.
For instance, cold air can be used to block the radially inward progression of the hot gases. Cold air from the compressor section (not shown) of the engine can be provided to the radially inner region 30 to cool the components and to physically impede the progress of the hot gases from the radially outer region 28 to the radially inner region 30 of the turbine 10. In addition, the cold air can mix with the hot gases to reduce the temperature of the gases to a mixing temperature. In addition, the discs 16 can be shielded from the hot gases by a cover plate 32, also known as a ring segment, that is secured to the disc 16. The cover plate 32 can cover at least a portion of the disc 16. A cover plate 32 can be provided on the axial upstream face 34 of the disc 16 and/or on the axial downstream face 36 of the disc 16.
Another method of reducing hot gas flow into the radially inner region 30 of the turbine 10 is to make a tortuous flow path, such as by providing a labyrinth-type sealing system in the axial gaps 26. To that end, the cover plate 32 can provide one or more axially extending arms 38. Each arm 38 can have a sealing surface 40, as shown in FIG. 2. Similarly, the neighboring stationary component, such as the pre-swirler 24, can have a plurality of axially extending protrusions 42. Each protrusion 42 can have a sealing surface 44. The sealing surfaces 40 of the arms 38 and the sealing surfaces 44 of the protrusions 42 are spaced from and substantially parallel to each other to form an annular gap 46 therebetween. The sealing surfaces 40, 44 are substantially parallel to the longitudinal axis 14 of the rotor.
While it is preferred if the gap 46 between the sealing surfaces 46, 50 is as small as possible, the gap 46 cannot be entirely eliminated because, during transient conditions, such as engine startup or part load operation, the rotating parts (blades 18, rotor 12, and discs 16) and the stationary parts (blade rings, vanes 20, and components attached to the vane) thermally expand at different rates. Thus, the gap 46 between the sealing surfaces 40, 44 is based on the cold condition with an understanding of the thermal behavior of the turbine components during engine operation. Under some operating conditions, particularly at steady state, the gap 46 between the sealing surfaces 40, 44 can increase. The consequences of such an increase in the size of the gap 46 can vary depending on the location in the turbine. In some instances, a larger gap can result in a greater mass flow of hot gases into the radial inner region 30 of the turbine 10, thereby requiring additional cooling air to be supplied for purposes of blocking. In other instances, the mass flow of cooling air leaking into the radial outer region 28 of the turbine 10 may increase, thereby causing performance losses. In either case, there can be a decrease in the output and efficiency of the engine.
Because the gap 46 is formed by surfaces that are substantially parallel to the longitudinal axis 14 of the rotor 12, the size of the gap 46 can only be adjusted by radial movement of the cover plate 32 and the components operatively connected thereto or by radial movement of the vane 20 or any component attached to the vane 20, such as the pre-swirler 24. Achieving such radial movement is difficult during engine operation. Thus, there is a need for a system that allows for greater flexibility in controlling the size of such leakage gaps.